CN105651287A - Solar flare TDOA (Time Difference Of Arrival) measurement, integrated navigation method and integrated navigation system - Google Patents

Abstract

The invention provides solar flare TDOA (Time Difference Of Arrival) measurement, anintegrated navigation method and an integrated navigation system, which belong to the field of spacecraft autonomous navigation. The solar flare TDOA measurement comprises the following steps of acquiring flare arrival time directly from the sun and flare arrival time reflected by the Mars; calculating a difference value between the two flare arrival times; building a solar flare TDOA measurement model. The integrated navigation method comprises the following steps of building a direction-findingnavigation model, building the solar flare TDOA measurement model, utilizing a filter to filter, and adopting the direction-finding model when the solar flare TDOA is not acquired. The solar flare TDOA measurement, and theintegrated navigation method and theintegrated navigationsystem provided by the invention solve the problem of large radial deviation of direction- finding navigation, are high in positioning precision, and have lower requirement on an instrument and ephemeris. Therefore, the solar flare TDOA measurement, and the integrated navigation method and the integrated navigation system provided by the invention has the important practical significance on spacecraftautonomous navigation.

Description

A kind of solar flare difference time of arrival is measured and Combinated navigation method, system

Technical field

The invention belongs to Spacecraft Autonomous Navigation field, in particular to a kind of solar flare difference time of arrival measuring technology scheme and the Spacecraft Autonomous Navigation technical scheme based on solar flare information time of arrival.

Background technology

Navigation information is most important for the success or failure of survey of deep space. By the impact of remote and long time delay, terrestrial station cannot provide the navigation information of real-time high-precision, particularly in the section of catching. And Spacecraft Autonomous Navigation system can accomplish this point. Therefore, for catching Duan Eryan, celestial autonomous navigation is extremely important, and particularly spacecraft is relative to the position of target celestial body, speed information.

At present, in deep-space detection field, there is following several autonomous navigation method: (1) X-ray pulsar range-finding navigation. X-ray pulsar navigation can provide precision distance measurement information. But, spacecraft is more important than absolute location relative to the position of target celestial body. If target celestial body star is gone through relatively big error occurs, also correspondingly there is relatively big error relative to the position of target celestial body in spacecraft. (2) direction finding navigation. Direction finding navigation is traditional star-navigation mode, obtains the azimuth information of spacecraft relative to nearly celestial body by measuring nearly celestial body. But, the method cannot provide range information between the spacecraft of high precision and nearly celestial body. (3) test the speed navigation. Test the speed the spectrum frequency displacement navigating through and measuring fixed star to obtain the speed information of spacecraft relative to fixed star. The navigation method that tests the speed cannot directly provide positional information. Positional information is obtained by integrating rate information, therefore must there is bigger integration error.

In sum, when target celestial body star go through there is error, X-ray pulsar navigation and navigation of testing the speed cannot provide the navigation information relative to target celestial body of high precision. Though direction finding navigation does not affect by ephemeris error, but precision is extremely low diametrically.

Summary of the invention

The present invention proposes a kind of solar flare difference time of arrival (TimeDifferenceOfArrival, TDOA) measuring technology scheme, it is intended that spacecraft provides the range-finding navigation information of high precision. On this basis, it is combined by the present invention with traditional direction finding navigation, it is proposed to a kind of solar flare TDOA/ direction finding integrated navigation technical scheme, it is intended to the survey of deep space section of catching is for spacecraft offer is real-time, the independent navigation information of high precision.

The present invention provides a kind of solar flare difference time of arrival measuring method, utilizes the spacecraft detected by solar system inner planet to realize measuring, spacecraft arranges the first photodetector and the 2nd photodetector, and measuring process comprises the following steps,

Steps A 1, adopts the first photodetector on spacecraft to record the time that solar flare directly arrives spacecraft, according to spacecraft at t₁Position r (the t in moment₁), calculate the time t that solar flare photon leaves sun barycenter₀It is as follows,

c��(t₁-t₀)=| r (t₁)|

Wherein, c is the light velocity;

Steps A 2, utilizes planet star to go through, and calculates the moment t that solar flare photon arrives planet₂Planetary position r now_M(t₂) as follows,

c��(t₂-t₀)=| r_M(t₂)|

Steps A 3, adopts the 2nd photodetector on spacecraft to record the time t recording and arriving spacecraft through planet reflection, obtains solar flare difference t-t time of arrival₁; Tentatively set up solar flare TDOA model as follows,

c��(t-t₁)=| r (t)-r_M(t₂)|-|r(t₁)|+|r_M(t₂)|+��

Wherein, �� is TDOA measuring error, and TDOA represents difference time of arrival, and spacecraft is r (t) in the position in moment t;

Steps A 4, carries out Geometric corrections to steps A 3 gained solar flare TDOA model, obtains accurate model, measures according to gained model realization difference time of arrival; Carry out Geometric corrections and comprise following sub-step,

Steps A 41, for the reflection spot not star catalogue face of being expert at planet barycenter, calculates angle alpha+beta as follows,

|r-r_M|²+|r_M|²-|r²=2 | r-r_M|��|r_M|cos(��+��)

Wherein, r, r_MAnd r_M' M represents that spacecraft is in position r (t) in moment t, at moment t respectively₂Corresponding Mars centroid position r_M(t₂) and reflection spot position r_M��(t₂), vector r-r_MAnd r_M��-r_MAngle be ��, vector-r_MAnd r_M��M-r_MAngle be ��;

Carry out initialize as follows,

��=��=(alpha+beta)/2

Steps A 42, if vector r_M-r and r_MThe angle of '-r is �� ', vector r_MAnd r_MThe angle of ' M is �� ', calculates �� ' and �� ' as follows,

${\left|r_M^{\′}\right|}^2=R_M^2+{\left|r_M\right|}^2-2R_M\·\left|r_M\right|cos\left(\mathrm{\β}\right)$

$R_M^2={\left|r_M^{\′}\right|}^2+{\left|r_M\right|}^2-2\left|r_M^{\′}\right|\·\left|r_M\right|cos\left({\mathrm{\β}}^{\′}\right)$

${\left|r_M^{\′}-r\right|}^2=R_M^2+{\left|r_M-r\right|}^2-2R_M\·\left|r_M-r\right|cos\left(\mathrm{\α}\right)$

$R_M^2={\left|r_M^{\′}-r\right|}^2+{\left|r_M-r\right|}^2-2\left|r_M^{\′}-r\right|\·\left|r_M-r\right|\·cos\left({\mathrm{\α}}^{\′}\right)$

Wherein, R_MFor planet radius;

Steps A 43, �� and �� is as follows in adjustment,

��=��-[(��+�� ')-(��+�� ')]/2

��=��+[(��+�� ')-(��+�� ')]/2

Steps A 44, returns steps A 42, until following formula is set up, now | and r_M' | with | r_M'-r | it is required distance;

��+�� '=��+�� '=��

Steps A 45, solar flare TDOA model steps A 3 tentatively set up is revised, and obtains correction model as follows,

t-t₁₌(r_M��-r|+|r_M��-|r-v��(t-t₁))/c+��

Wherein, v is the speed of spacecraft.

And, described planet is Mars, and described spacecraft is Mars detector.

A kind of solar flare difference time of arrival measuring system of the corresponding offer of the present invention, for utilizing the spacecraft detected by solar system inner planet to realize measuring, spacecraft arranges the first photodetector and the 2nd photodetector, and arranges with lower module,

First module, for adopting the first photodetector on spacecraft to record the time that solar flare directly arrives spacecraft, according to spacecraft at t₁Position r (the t in moment₁), calculate the time t that solar flare photon leaves sun barycenter₀It is as follows,

c��(t₁-t₀)=| r (t₁)|

Wherein, c is the light velocity;

2nd module, for utilizing planet star to go through, calculates the moment t that solar flare photon arrives planet₂Planetary position r now_M(t₂) as follows,

c��(t₂-t₀)=| r_M(t₂)|

3rd module, records the time t recording and arriving spacecraft through planet reflection, obtains solar flare difference t-t time of arrival for adopting the 2nd photodetector on spacecraft₁; Tentatively set up solar flare TDOA model as follows,

c��(t-t₁)=| r (t)-r_M(t₂)|-|r(t₁)|+|r_M(t₂)|+��

Wherein, �� is TDOA measuring error, and TDOA represents difference time of arrival, and spacecraft is r (t) in the position in moment t;

Four module, for the 3rd module gained solar flare TDOA model is carried out Geometric corrections, obtains accurate model, measures according to gained model realization difference time of arrival; Comprise following submodule block,

First submodule block, for for the reflection spot not star catalogue face of being expert at planet barycenter, calculating angle alpha+beta as follows,

|r-r_M|²+|r_M|²-|r|²=2 | r-r_M|��|r_M|cos(��+��)

Wherein, r, r_MAnd r_M' represent that spacecraft is in position r (t) in moment t, at moment t respectively₂Corresponding Mars centroid position r_M(t₂) and reflection spot position r_M��(t₂), vector r-r_MAnd r_M��-r_MAngle be ��, vector-r_MAnd r_M��-r_MAngle be ��;

Carry out initialize as follows,

��=��=(alpha+beta)/2

2nd submodule block, for establishing vector r_M-r and r_MThe angle of '-r is �� ', vector r_MAnd r_M' angle be �� ', calculate �� ' and �� ' as follows,

${\left|r_M^{\′}\right|}^2=R_M^2+{\left|r_M\right|}^2-2R_M\·\left|r_M\right|cos\left(\mathrm{\β}\right)$

$R_M^2={\left|r_M^{\′}\right|}^2+{\left|r_M\right|}^2-2\left|r_M^{\′}\right|\·\left|r_M\right|cos\left({\mathrm{\β}}^{\′}\right)$

${\left|r_M^{\′}-r\right|}^2=R_M^2+{\left|r_M-r\right|}^2-2R_M\·\left|r_M-r\right|cos\left(\mathrm{\α}\right)$

$R_M^2={\left|r_M^{\′}-r\right|}^2+{\left|r_M-r\right|}^2-2\left|r_M^{\′}-r\right|\·\left|r_M-r\right|\·cos\left({\mathrm{\α}}^{\′}\right)$

Wherein, R_MFor planet radius;

3rd submodule block, as follows for adjusting �� and ��,

��=��-[(��+�� ')-(��+�� ')]/2

��=��+[(��+�� ')-(��+�� ')]/2

4th submodule block, for ordering the 2nd sub-module work, until following formula is set up, now | r_M' | with | r_M'-r | it is required distance;

��+�� '=��+�� '=��

5th submodule block, revises for the solar flare TDOA model the 3rd module tentatively set up, obtains correction model as follows,

t-t₁=(r_M��-r|+|r_M��|-|r-v��(t-t₁))/c+��

Wherein, v is the speed of spacecraft.

And, described planet is Mars, and described spacecraft is Mars detector.

The present invention also provides a kind of Combinated navigation method measured based on solar flare difference time of arrival, comprises the following steps,

Step 1, the track kinetic model setting up spacecraft is as follows,

$\stackrel{\·}{X}\left(T\right)=f(X,T)+\mathrm{\ω}\left(T\right)$

Wherein,It is the derivative of the state vector X of spacecraft,For moment TThe state metastasis model that f (X, T) is spacecraft, �� (T) is the navigationsystem noise of moment T spacecraft;

Step 2, sets up direction finding model as follows,

$Z=\frac{R-R_M}{\left|R-R_M\right|}+\mathrm{\υ}$

Wherein, �� is direction finding noise, and Z is the orientation vector of spacecraft relative to planet, R, R_MRepresent spacecraft respectively

Position, Mars centroid position;

Step 3, sets up solar flare TDOA model as follows,

t-t₁=(r_M��M-r|+|r_M��|-|r-v��(t-t₁))/c+��

Wherein, t-t₁For solar flare difference time of arrival, t₁For solar flare directly arrives the time of spacecraft, t is the time arriving spacecraft through planet reflection, r and r_M' respectively represent spacecraft in position r (t) in moment t, solar flare photon arrive Mars moment t₂Corresponding reflection spot position r_M��(t₂), v is the speed of spacecraft, and c is the light velocity, and �� is TDOA measuring error;

Step 4, when not obtaining solar flare TDOA, selects step 2 gained direction finding model, carries out filtering according to gained track kinetic model in step 1; When obtaining solar flare TDOA, select step 3 gained TDOA model, carry out filtering according to gained track kinetic model in step 1; Filtering obtains navigation information.

And, described filtering adopts spreading kalman wave filter to realize.

A kind of integrated navigation system measured based on solar flare difference time of arrival of the corresponding offer of the present invention, comprises with lower unit,

Unit first, as follows for setting up the track kinetic model of spacecraft,

$\stackrel{\·}{X}\left(T\right)=f(X,T)+\mathrm{\ω}\left(T\right)$

Wherein,It is the derivative of the state vector X of spacecraft,For moment TThe state metastasis model that f (X, T) is spacecraft, �� (T) is the navigationsystem noise of moment T spacecraft;

Second unit, as follows for setting up direction finding model,

$Z=\frac{R-R_M}{\left|R-R_M\right|}+\mathrm{\υ}$

Wherein, �� is direction finding noise, and Z is the orientation vector of spacecraft relative to planet, R, R_MRepresent position, the Mars centroid position of spacecraft respectively;

Unit the 3rd, as follows for setting up solar flare TDOA model,

t-t₁=(r_M��-r|+|r_M��|-|r-v��(t-t₁))/c+��

Wherein, t-t₁For solar flare difference time of arrival, t₁For solar flare directly arrives the time of spacecraft, t is the time arriving spacecraft through planet reflection, r and r_M' respectively represent spacecraft in position r (t) in moment t, solar flare photon arrive Mars moment t₂Corresponding reflection spot position r_M��(t₂), v is the speed of spacecraft, and c is the light velocity, and �� is TDOA measuring error;

Unit the 4th, during for not obtaining solar flare TDOA, selects second unit gained direction finding model, carries out filtering according to gained track kinetic model in Unit first; When obtaining solar flare TDOA, select the 3rd unit gained TDOA model, carry out filtering according to gained track kinetic model in Unit first; Filtering obtains navigation information.

And, described filtering adopts spreading kalman wave filter to realize.

The present invention's advantage compared with prior art is:

(1) Spacecraft Autonomous Navigation system needs detector and relevant chronometer data. Existing method all needs to develop new detector before realization, and carries out long-term sky patrol and obtain chronometer data. And the present invention only needs photodetector and target celestial body star to go through. This equipment and chronometer data are all existing, it is not necessary to development or collection again, have saved cost and time.

(2) the present invention can go through at the target celestial body star of low speed photodetector, low precision, without normal operation when solar flare timing model. Therefore, the present invention equipment and star are gone through require low, be easy to realize.

(3) navigation observed quantity of the present invention is the azimuth-range information of spacecraft relative to target celestial body. Therefore, the method is not by the impact of target celestial body ephemeris error, it is provided that the navigation information relative to target celestial body of high precision.

Accompanying drawing explanation

Fig. 1 is the solar flare difference time of arrival measuring principle schematic diagram of the embodiment of the present invention.

Fig. 2 is the Geometric corrections schematic diagram of the solar flare TDOA measurement of the embodiment of the present invention.

Embodiment

Technical solution of the present invention can adopt computer software mode to support automatic operational scheme. Below in conjunction with drawings and Examples, technical solution of the present invention is described in detail.

Survey of deep space device is commonly referred to as spacecraft. The spacecraft of Mars of flying to can be called Mars detector.

The present invention is taking Mars detector as embodiment, and the detector for other solar system inner planets also can adopt same method to realize.

First Mars detector track is provided, as shown in table 1.

The initial orbit parameter of table 1 Mars detector

The ultimate principle of solar flare TDOA is as shown in Figure 1. Path D is the path that solar flare photon directly flies to spacecraft from the sun, and path M is that solar photon flies to the path of spacecraft again through Mars reflection. Mars detector is provided with two photodetectors.

A kind of solar flare difference time of arrival measuring method that embodiment provides, is specially:

Steps A 1: obtain the time t that solar flare directly arrives spacecraft₁. This value can be recorded by the photodetector (being designated as the first photodetector) on spacecraft. According to spacecraft at t₁Position r (the t in moment₁) and formula (1), the time t that solar flare photon leaves sun barycenter can be calculated₀��

c��(t₁-t₀)=| r (t₁) | (1) wherein, c is the light velocity. | | represent vector delivery, lower same.

Steps A 2: utilize Mars star to go through and formula (2), can calculate the moment t that solar flare photon arrives Mars₂Mars position r now_M(t₂)��

c��(t₂-t₀)=| r_M(t₂)|(2)

Steps A 3: record the time t arriving spacecraft through Mars reflection.This value can be recorded by another photodetector (being designated as the 2nd photodetector) on spacecraft. t-t₁It is exactly solar flare TDOA. Solar flare TDOA model can tentatively be set up, shown in (3).

c��(t-t₁)=| r (t)-r_M(t₂)|-|r(t₁)|+|r_M(t₂)|+��(3)

Wherein, �� is TDOA measuring error. Spacecraft is r (t) in the position of t, and this value is unknown quantity, is the state that navigationsystem needs to estimate.

Steps A 4: this model is carried out Geometric corrections, obtains accurate model, measures according to gained model realization difference time of arrival.

Solar flare TDOA Geometric corrections principle is as shown in Figure 2. Reflection spot not at Mars barycenter, and Mars surface. For this problem, the present invention carries out Geometric corrections.

Steps A 41: according to formula (4), calculates the value of alpha+beta. And with formula (5) initialize �� and ��.

|r-r_M|²+|r_M|²-|r|²=2 | r-r_M|��|r_M| cos (alpha+beta) (4) wherein, r, r_MAnd r_M' represent r (t), r respectively_M(t₂) and r_M��(t₂), t₂Definition be the same. Mars centroid position and reflection spot position are respectively r_MAnd r_M', i.e. r_M(t₂) represent that solar flare photon arrives the moment t of Mars₂Corresponding Mars centroid position, r_M��(t₂) represent that solar flare photon arrives the moment t of Mars₂Corresponding reflection spot position. Vector r-r_MAnd r_M��-r_MAngle be ��, vector-r_MAnd r_M��-r_MAngle be ��. First original hypothesis is as follows:

��=��=(alpha+beta)/2 (5)

Steps A 42: utilize formula (6-9) to calculate �� ' and �� '. Vector r_M-r and r_MThe angle of '-r is �� ', vector r_MAnd r_M' angle be �� '. R_MFor Mars radius.

${\left|r_M^{\′}\right|}^2=R_M^2+{\left|r_M\right|}^2-2R_M\·\left|r_M\right|cos\left(\mathrm{\β}\right)---\left(6\right)$

$R_M^2={\left|r_M^{\′}\right|}^2+{\left|r_M\right|}^2-2\left|r_M^{\′}\right|\·\left|r_M\right|cos\left({\mathrm{\β}}^{\′}\right)---\left(7\right)$

${\left|r_M^{\′}-r\right|}^2=R_M^2+{\left|r_M-r\right|}^2-2R_M\·\left|r_M-r\right|cos\left(\mathrm{\α}\right)---\left(8\right)$

$R_M^2={\left|r_M^{\′}-r\right|}^2+{\left|r_M-r\right|}^2-2\left|r_M^{\′}-r\right|\·\left|r_M-r\right|\·cos\left({\mathrm{\α}}^{\′}\right)---\left(9\right)$

Steps A 43: utilize formula (10-11), adjustment �� and ��.

��=��-[(��+�� ')-(��+�� ')]/2 (10)

��=��+[(��+�� ')-(��+�� ')]/2 (11)

Steps A 44: repeating step A42-A44, until formula (12) is set up. Now | r_M' |, | r_M'-r | it is required distance.

��+�� '=��+�� '=�� (12)

Steps A 45: formula (3) revised, obtains correction model such as formula shown in (13).

t-t₁=(r_M��-r|+|r_M��|-|r-v��(t-t₁))/c+��(13)

Wherein, v is the speed of spacecraft.

Embodiment is based on above solar flare TDOA method, it is provided that a kind of solar flare TDOA/ direction finding Combinated navigation method, be specially:

Step B1: the track kinetic model setting up spacecraft

The track kinetic model setting up spacecraft described in step B1, its specific implementation process is:

Because the state vector X of spacecraft is:

$X=\left[\begin{array}{c}r\\ v\end{array}\right]---\left(14\right)$

Wherein, r=[x, y, z]^TWith v=[v_x,v_y,v_z]^TPosition and the velocity vector being respectively spacecraft, the component of the position that x, y, z are respectively spacecraft on three axles, v_x,v_y,v_zThe component of the speed being respectively spacecraft on three axles;

Then the track kinetic model of spacecraft is:

$\left\{\begin{array}{c}\stackrel{\·}{x}=v_x\\ \stackrel{\·}{y}=v_x\\ \stackrel{\·}{z}=v_z\\ {\stackrel{\·}{v}}_x=-{\mathrm{\μ}}_s\frac{x}{r_{ps}^3}-{\mathrm{\μ}}_m\[\frac{x-x_1}{r_{pm}^3}+\frac{x_1}{r_{sm}^3}\]-{\mathrm{\μ}}_e\[\frac{x-x_2}{r_{pe}^3}+\frac{x_2}{r_{se}^3}\]+{\mathrm{\ΔF}}_x\\ {\stackrel{\·}{v}}_y=-{\mathrm{\μ}}_s\frac{y}{r_{ps}^3}-{\mathrm{\μ}}_m\[\frac{y-y_1}{r_{pm}^3}+\frac{y_1}{r_{sm}^3}\]-{\mathrm{\μ}}_e\[\frac{y-y_2}{r_{pe}^3}+\frac{y_2}{r_{se}^3}\]+{\mathrm{\ΔF}}_y\\ {\stackrel{\·}{v}}_z=-{\mathrm{\μ}}_s\frac{z}{r_{ps}^3}-{\mathrm{\μ}}_m\[\frac{z-z_1}{r_{pm}^3}+\frac{z_1}{r_{sm}^3}\]-{\mathrm{\μ}}_e\[\frac{z-z_2}{r_{pe}^3}+\frac{z_2}{r_{se}^3}\]+{\mathrm{\ΔF}}_z\end{array}\right.---\left(15\right)$

Wherein,It is respectively x, y, z, v_x,v_y,v_zDerivative,

Formula (5) can represent:

$\stackrel{\·}{X}\left(T\right)=f(X,T)+\mathrm{\ω}\left(T\right)---\left(16\right)$

Wherein,It is the derivative of X,For moment TThe state metastasis model that f (X, T) is spacecraft, [x₁,y₁,z₁] and [x₂,y₂,z₂] it is Mars and the earth Relative position vector relative to solar system barycenter respectively, ��_s,��_m,��_eIt is the gravity constant of the sun, Mars and the earth respectively, r_ps,r_pm,r_peBe respectively spacecraft to the distance between sun barycenter, Mars barycenter and earth centroid, its calculation formula is: $r_{ps}=\sqrt{x^2+y^2+z^2},$ $r_{pm}=\sqrt{{(x-x_1)}^2+{(y-y_1)}^2+{(z-z_1)}^2},$ $r_{pe}=\sqrt{{(x-x_2)}^2+{(y-y_2)}^2+{(z-z_2)}^2};$ $r_{sm}=\sqrt{x_1^2+y_1^2+z_1^2},$ $r_{se}=\sqrt{x_2^2+y_2^2+z_2^2}$ It is that Mars barycenter, earth centroid divide the distance being clipped between sun barycenter respectively; Navigationsystem noise ��=[0,0,0, the �� F of spacecraft_x,��F_y,��F_z]^T, wherein, �� F_x����F_yWith �� F_zBeing the component of disturbing force on three axles, �� (T) is the navigationsystem noise of moment T spacecraft.

Step B2: set up direction finding model.

$Z=\frac{R-R_M}{\left|R-R_M\right|}+\mathrm{\υ}---\left(17\right)$

Wherein, �� is direction finding noise. Z is the orientation vector of spacecraft relative to Mars. Now, R, R_MRepresent position, the Mars centroid position of spacecraft respectively.

Step B3: set up solar flare TDOA model. Refer to above-mentioned relative navigation speed-measuring method, adopt formula (13).

Step B4: when not obtaining solar flare TDOA, selects step B2 gained direction finding model, carries out filtering according to gained track kinetic model in step B1; When obtaining solar flare TDOA, select step B3 gained TDOA model, carry out filtering according to gained track kinetic model in step B1, obtain navigation information.

Embodiment utilizes spreading kalman filter filtering. The state vector of wave filter is position and the velocity vector of spacecraft, and filtering result is exactly position and the speed of spacecraft. Measurement model system of selection in Navigation Filter is as follows:

When not obtaining solar flare TDOA, select step B2 gained direction finding model. When obtaining solar flare TDOA, select step B3 gained TDOA model.

State metastasis model in Navigation Filter is the track kinetic model shown in B1 step Chinese style (16).

When specifically implementing, adopt the implementation of other wave filters similar.

Wave filter parameter is as shown in table 2:

Table 2 Navigation Filter parameter

Wherein, P (0) for original state error matrix, Q be state-noise covariance,I.e. q₁Square,I.e. q₂Square.

When specifically implementing, above flow process can adopt computer software technology to realize automatic operational scheme, it is possible to adopts software modularity technology to realize corresponding system.

A kind of solar flare difference time of arrival measuring system of the corresponding offer of the present invention, for utilizing the spacecraft detected by solar system inner planet to realize measuring, spacecraft arranges the first photodetector and the 2nd photodetector, and arranges with lower module,

First module, for adopting the first photodetector on spacecraft to record the time that solar flare directly arrives spacecraft, according to spacecraft at t₁Position r (the t in moment₁), calculate the time t that solar flare photon leaves sun barycenter₀It is as follows,

c��(t₁-t₀)=| r (t₁)|

Wherein, c is the light velocity;

2nd module, for utilizing planet star to go through, calculates the moment t that solar flare photon arrives planet₂Planetary position r now_M(t₂) as follows,

c��(t₂-t₀)=| r_M(t₂)|

3rd module, records the time t recording and arriving spacecraft through planet reflection, obtains solar flare difference t-t time of arrival for adopting the 2nd photodetector on spacecraft₁; Tentatively set up solar flare TDOA model as follows,

c��(t-t₁)=| r (t)-r_M(t₂)|-|r(t₁)|+|r_M(t₂)|+��

Wherein, �� is TDOA measuring error, and TDOA represents difference time of arrival, and spacecraft is r (t) in the position in moment t;

Four module, for the 3rd module gained solar flare TDOA model is carried out Geometric corrections, obtains accurate model, measures according to gained model realization difference time of arrival; Comprise following submodule block,

First submodule block, for for the reflection spot not star catalogue face of being expert at planet barycenter, calculating angle alpha+beta as follows,

|r-r_M|²+|r_M|²-|r²=2 | r-r_M|��|r_M|cos(��+��)

Wherein, r, r_MAnd r_M' represent that spacecraft is in position r (t) in moment t, at moment t respectively₂Corresponding Mars centroid position r_M(t₂) and reflection spot position r_M��(t₂), vector r-r_MAnd r_M��-r_MAngle be ��, vector-r_MAnd r_M��-r_MAngle be ��;

Carry out initialize as follows,

��=��=(alpha+beta)/2

2nd submodule block, for establishing vector r_M-r and r_MThe angle of '-r is �� ', vector r_MAnd r_M' angle be �� ', calculate �� ' and �� ' as follows,

${\left|r_M^{\′}\right|}^2=R_M^2+{\left|r_M\right|}^2-2R_M\·\left|r_M\right|cos\left(\mathrm{\β}\right)$

$R_M^2={\left|r_M^{\′}\right|}^2+{\left|r_M\right|}^2-2\left|r_M^{\′}\right|\·\left|r_M\right|cos\left({\mathrm{\β}}^{\′}\right)$

${\left|r_M^{\′}-r\right|}^2=R_M^2+{\left|r_M-r\right|}^2-2R_M\·\left|r_M-r\right|cos\left(\mathrm{\α}\right)$

$R_M^2={\left|r_M^{\′}-r\right|}^2+{\left|r_M-r\right|}^2-2\left|r_M^{\′}-r\right|\·\left|r_M-r\right|\·cos\left({\mathrm{\α}}^{\′}\right)$

Wherein, R_MFor planet radius;

3rd submodule block, as follows for adjusting �� and ��,

��=��-[(��+�� ')-(��+�� ')]/2

��=��+[(��+�� ')-(��+�� ')]/2

4th submodule block, for ordering the 2nd sub-module work, until following formula is set up, now | r_M' | with | r_M'-r | it is required distance;

��+�� '=��+�� '=��

5th submodule block, revises for the solar flare TDOA model the 3rd module tentatively set up, obtains correction model as follows,

t-t₁=(r_M��-r|+|r_M��|-|r-v��(t-t₁))/c+��

Wherein, v is the speed of spacecraft.

A kind of integrated navigation system measured based on solar flare difference time of arrival of the corresponding offer of the present invention, comprises with lower unit,

Unit first, as follows for setting up the track kinetic model of spacecraft,

$\stackrel{\·}{X}\left(T\right)=f(X,T)+\mathrm{\ω}\left(T\right)$

Wherein,It is the derivative of the state vector X of spacecraft,For moment TThe state metastasis model that f (X, T) is spacecraft, �� (T) is the navigationsystem noise of moment T spacecraft;

Second unit, as follows for setting up direction finding model,

$Z=\frac{R-R_M}{\left|R-R_M\right|}+\mathrm{\υ}$

Wherein, �� is direction finding noise, and Z is the orientation vector of spacecraft relative to planet, R, R_MRepresent position, the Mars centroid position of spacecraft respectively;

Unit the 3rd, as follows for setting up solar flare TDOA model,

t-t₁=(| r_M��-r|+r_M��|-r-v��(t-t₁))/c+��

Wherein, t-t₁For solar flare difference time of arrival, t₁For solar flare directly arrives the time of spacecraft, t is the time arriving spacecraft through planet reflection, r and r_M' respectively represent spacecraft in position r (t) in moment t, solar flare photon arrive Mars moment t₂Corresponding reflection spot position r_M��(t₂), v is the speed of spacecraft, and c is the light velocity, and �� is TDOA measuring error;

Unit the 4th, during for not obtaining solar flare TDOA, selects second unit gained direction finding model, carries out filtering according to gained track kinetic model in Unit first; When obtaining solar flare TDOA, select the 3rd unit gained TDOA model, carry out filtering according to gained track kinetic model in Unit first; Filtering obtains navigation information.

Each module or unit realize can see method respective description, and it will not go into details in the present invention.

Specific embodiment described herein is only to the present invention's spirit explanation for example. Described specific embodiment can be made various amendment or supplements or adopt similar mode to substitute by those skilled in the art, but can't deviate the spirit of the present invention or surmount the scope that appended claims defines.

Claims (8)

1. a solar flare difference time of arrival measuring method, it is characterised in that: utilizing the spacecraft detected by solar system inner planet to realize measuring, spacecraft arranges the first photodetector and the 2nd photodetector, measuring process comprises the following steps,

Steps A 1, adopts the first photodetector on spacecraft to record the time that solar flare directly arrives spacecraft, according to spacecraft at t₁Position r (the t in moment₁), calculate the time t that solar flare photon leaves sun barycenter₀It is as follows,

c��(t₁-t₀)=| r (t₁)|

Wherein, c is the light velocity;

Steps A 2, utilizes planet star to go through, and calculates the moment t that solar flare photon arrives planet₂Planetary position r now_M(t₂) as follows,

c��(t₂-t₀)=| r_M(t₂)|

Steps A 3, adopts the 2nd photodetector on spacecraft to record the time t recording and arriving spacecraft through planet reflection, obtains solar flare difference t-t time of arrival₁; Tentatively set up solar flare TDOA model as follows,

c��(t-t₁)=| r (t)-r_M(t₂)|-|r(t₁)|+|r_M(t₂)|+��

Wherein, �� is TDOA measuring error, and TDOA represents difference time of arrival, and spacecraft is r (t) in the position in moment t.

Steps A 4, carries out Geometric corrections to steps A 3 gained solar flare TDOA model, obtains accurate model, measures according to gained model realization difference time of arrival; Carry out Geometric corrections and comprise following sub-step,

Steps A 41, for the reflection spot not star catalogue face of being expert at planet barycenter, calculates angle alpha+beta as follows,

r-r_M|²+|r_M|²-|r²=2 | r-r_M|��|r_M|cos(��+��)

Wherein, r, r_MAnd r_M' represent that spacecraft is in position r (t) in moment t, at moment t respectively₂Corresponding Mars centroid position r_M(t₂) and reflection spot position r_M��(t₂), vector r-r_MAnd r_M��-r_MAngle be ��, vector-r_MAnd r_M��-r_MAngle be ��;

Carry out initialize as follows,

��=��=(alpha+beta)/2

Steps A 42, if vector r_M-r and r_MThe angle of '-r is �� ', vector r_MAnd r_M' angle be �� ', calculate �� ' and �� ' as follows,

${\left|r_M^{\′}\right|}^2=R_M^2+{\left|r_M\right|}^2-2R_M\·\left|r_M\right|cos\left(\mathrm{\β}\right)$

$R_M^2={\left|r_M^{\′}\right|}^2+{\left|r_M\right|}^2-2\left|r_M^{\′}\right|\·\left|r_M\right|cos\left({\mathrm{\β}}^{\′}\right)$

${\left|r_M^{\′}-r\right|}^2=R_M^2+{\left|r_M-r\right|}^2-2R_M\·\left|r_M-r\right|cos\left(\mathrm{\α}\right)$

$R_M^2={\left|r_M^{\′}-r\right|}^2+{\left|r_M-r\right|}^2-2\left|r_M^{\′}-r\right|\·\left|r_M-r\right|\·cos\left({\mathrm{\α}}^{\′}\right)$

Wherein, R_MFor planet radius;

Steps A 43, �� and �� is as follows in adjustment,

��=��-[(��+�� ')-(��+�� ')]/2

��=��+[(��+�� ')-(��+�� ')]/2

Steps A 44, returns steps A 42, until following formula is set up, now | and r_M' | with | r_M'-r | it is required distance;

��+�� '=��+�� '=��

Steps A 45, solar flare TDOA model steps A 3 tentatively set up is revised, and obtains correction model as follows,

t-t₁=(r_M��-r|+|r_M��|-|r-v��(t-t₁))/c+��

Wherein, v is the speed of spacecraft.

2. solar flare difference time of arrival measuring method according to claim 1, it is characterised in that: described planet is Mars, and described spacecraft is Mars detector.

3. a solar flare difference time of arrival measuring system, it is characterised in that: for utilizing the spacecraft detected by solar system inner planet to realize measuring, spacecraft arranges the first photodetector and the 2nd photodetector, and arranges with lower module,

First module, for adopting the first photodetector on spacecraft to record the time that solar flare directly arrives spacecraft, according to spacecraft at t₁Position r (the t in moment₁), calculate the time t that solar flare photon leaves sun barycenter₀It is as follows,

c��(t₁-t₀)=| r (t₁)|

Wherein, c is the light velocity;

2nd module, for utilizing planet star to go through, calculates the moment t that solar flare photon arrives planet₂Planetary position r now_M(t₂) as follows,

c��(t₂-t₀)=| r_M(t₂)|

3rd module, records the time t recording and arriving spacecraft through planet reflection, obtains solar flare difference t-t time of arrival for adopting the 2nd photodetector on spacecraft₁; Tentatively set up solar flare TDOA model as follows,

c��(t-t₁)=| r (t)-r_M(t₂)|-|r(t₁)|+|r_M(t₂)|+��

Wherein, �� is TDOA measuring error, and TDOA represents difference time of arrival, and spacecraft is r (t) in the position in moment t;

Four module, for the 3rd module gained solar flare TDOA model is carried out Geometric corrections, obtains accurate model, measures according to gained model realization difference time of arrival; Comprise following submodule block,

First submodule block, for for the reflection spot not star catalogue face of being expert at planet barycenter, calculating angle alpha+beta as follows,

|r-r_M|²+|r_M|²-|r|²=2 | r-r_M|��|r_M|cos(��+��)

Wherein, r, r_MAnd r_M' represent that spacecraft is in position r (t) in moment t, at moment t respectively₂Corresponding Mars centroid position r_M(t₂) and reflection spot position r_M��(t₂), vector r-r_MAnd r_M��-r_MAngle be ��, vector-r_MAnd r_M��-r_MAngle be ��;

Carry out initialize as follows,

��=��=(alpha+beta)/2

2nd submodule block, for establishing vector r_M-r and r_MThe angle of '-r is �� ', vector r_MAnd r_M' angle be �� ', calculate �� ' and �� ' as follows,

${\left|r_M^{\′}\right|}^2=R_M^2+{\left|r_M\right|}^2-2R_M\·\left|r_M\right|cos\left(\mathrm{\β}\right)$

$R_M^2={\left|r_M^{\′}\right|}^2+{\left|r_M\right|}^2-2\left|r_M^{\′}\right|\·\left|r_M\right|cos\left({\mathrm{\β}}^{\′}\right)$

${\left|r_M^{\′}-r\right|}^2=R_M^2+{\left|r_M-r\right|}^2-2R_M\·\left|r_M-r\right|cos\left(\mathrm{\α}\right)$

$R_M^2={\left|r_M^{\′}-r\right|}^2+{\left|r_M-r\right|}^2-2\left|r_M^{\′}-r\right|\·\left|r_M-r\right|\·cos\left({\mathrm{\α}}^{\′}\right)$

Wherein, R_MFor planet radius;

3rd submodule block, as follows for adjusting �� and ��,

��=��-[(��+�� ')-(��+�� ')]/2

��=��+[(��+�� ')-(��+�� ')]/2

4th submodule block, for ordering the 2nd sub-module work, until following formula is set up, now | r_M' | with | r_M'-r | it is required distance;

��+�� '=��+�� '=��

5th submodule block, revises for the solar flare TDOA model the 3rd module tentatively set up, obtains correction model as follows,

t-t₁=(r_M��-r|+|r_M��|-|r-v��(t-t₁))/c+��

Wherein, v is the speed of spacecraft.

4. solar flare difference time of arrival measuring system according to claim 3, it is characterised in that: described planet is Mars, and described spacecraft is Mars detector.

5. the Combinated navigation method measured based on solar flare difference time of arrival, it is characterised in that: comprise the following steps,

Step 1, the track kinetic model setting up spacecraft is as follows,

$\stackrel{\·}{X}\left(T\right)=f(X,T)+\mathrm{\ω}\left(T\right)$

Wherein,It is the derivative of the state vector X of spacecraft,For moment TThe state metastasis model that f (X, T) is spacecraft, �� (T) is the navigationsystem noise of moment T spacecraft;

Step 2, sets up direction finding model as follows,

$Z=\frac{R-R_M}{\left|R-R_M\right|}+\mathrm{\υ}$

Wherein, �� is direction finding noise, and Z is the orientation vector of spacecraft relative to planet, R, R_MRepresent position, the Mars centroid position of spacecraft respectively;

Step 3, sets up solar flare TDOA model as follows,

t-t₁=(| r_M��-r|+|r_M��|-|r-v��(t-t₁))/c+��

Wherein, t-t₁For solar flare difference time of arrival, t₁For solar flare directly arrives the time of spacecraft, t is the time arriving spacecraft through planet reflection, r and r_M' respectively represent spacecraft in position r (t) in moment t, solar flare photon arrive Mars moment t₂Corresponding reflection spot position r_M��(t₂), v is the speed of spacecraft, and c is the light velocity, and �� is TDOA measuring error;

Step 4, when not obtaining solar flare TDOA, selects step 2 gained direction finding model, carries out filtering according to gained track kinetic model in step 1; When obtaining solar flare TDOA, select step 3 gained TDOA model, carry out filtering according to gained track kinetic model in step 1; Filtering obtains navigation information.

6. the Combinated navigation method measured based on solar flare difference time of arrival according to claim 5, it is characterised in that: described filtering adopts spreading kalman wave filter to realize.

7. the integrated navigation system measured based on solar flare difference time of arrival, it is characterised in that: comprise with lower unit,

Unit first, as follows for setting up the track kinetic model of spacecraft,

$\stackrel{\·}{X}\left(T\right)=f(X,T)+\mathrm{\ω}\left(T\right)$

Wherein,It is the derivative of the state vector X of spacecraft,For moment TThe state metastasis model that f (X, T) is spacecraft, �� (T) is the navigationsystem noise of moment T spacecraft;

Second unit, as follows for setting up direction finding model,

$Z=\frac{R-R_M}{\left|R-R_M\right|}+\mathrm{\υ}$

Wherein, �� is direction finding noise, and Z is the orientation vector of spacecraft relative to planet, R, R_MRepresent spacecraft respectively

Position, Mars centroid position;

Unit the 3rd, as follows for setting up solar flare TDOA model,

t-t₁=(| r_M��-r|+|r_M��|-|r-v��(t-t₁))/c+��

Wherein, t-t₁For solar flare difference time of arrival, t₁For solar flare directly arrives the time of spacecraft, t is the time arriving spacecraft through planet reflection, r and r_M' respectively represent spacecraft in position r (t) in moment t, solar flare photon arrive Mars moment t₂Corresponding reflection spot position r_M��(t₂), v is the speed of spacecraft, and c is the light velocity, and �� is TDOA measuring error;

Unit the 4th, during for not obtaining solar flare TDOA, selects second unit gained direction finding model, carries out filtering according to gained track kinetic model in Unit first;When obtaining solar flare TDOA, select the 3rd unit gained TDOA model, carry out filtering according to gained track kinetic model in Unit first; Filtering obtains navigation information.

8. the Combinated navigation method measured based on solar flare difference time of arrival according to claim 7, it is characterised in that: described filtering adopts spreading kalman wave filter to realize.